Embodiments of the present invention relate to gas turbine engines and, more particularly, to methods and apparatus for providing cooling air to turbine airfoils within a gas turbine engine.
In a gas turbine engine, hot gas exits a combustor and is utilized by a turbine for conversion to mechanical energy. This mechanical energy drives an upstream high pressure compressor. The turbine comprises a plurality of rows of blades which are carried by a turbine rotor, alternating with rows of stationary nozzles. The turbine blades and nozzles are subjected to a flow of the corrosive, high-temperature combustion gases. These “hot section” components are typically cooled by a flow of relatively low-temperature coolant, such as air extracted (bled) from the compressor or compressor discharge air. Using air extracted from the cycle in this manner is chargeable to the thermodynamic cycle, increases specific fuel consumption (“SFC”), and is generally to be avoided or minimized whenever possible.
One known type of turbine cooling system uses inducers to collect compressor discharge air, accelerate it and turn it tangentially, and feed it to a turbine rotor. Typically, turbine cooling systems are physically configured to meet maximum cooling demand, as would be experienced during high-power operation such as takeoff or initial climb. This results in excess cooling capacity during other operating conditions such as cruise flight or descent. Cooling demand in these conditions is much lower and actually represents the majority of the time of engine operation.
Attempts have been made to modulate turbine cooling flow during other operating conditions. These typically require piping external to the engine case, which is subject to undesirable failure modes, or internally-actuated valves which are difficult to keep in good operating condition in the high-temperature environment within the engine.
Accordingly, there is a need for improved cooling systems which will provide cooling to an airfoil in a robust and economical manner.